A Gas Turbine Engine (GTE) can contain various different combinations of bladed rotors, such as axial compressor, radial or centrifugal compressor, axial turbine, radial-inflow turbine, and fan rotors. The thermal and mechanical demands placed on a particular bladed rotor can vary significantly across the rotor during GTE operation. Generally, the rotor blades are bathed in core gas flow and thus exposed to high temperature, chemically-harsh (e.g., corrosive and oxidative) environments. In contrast, the inner “hub disk” portion of the rotor is largely shielded from core gas flow, but may be subject to considerable mechanical stress resulting from the centrifugal forces acting on the rotor at high rotational speeds. Performance benefits can consequently be realized by fabricating the hub disk and rotor blades from different alloys tailored to their specific operating environments. For example, a so-called “inserted blade rotor” can be produced by attaching a number of bladed pieces composed of a first superalloy to a separately-fabricated rotor hub composed of a different superalloy. The bladed pieces are fabricated to include shanks, which are inserted into mating slots provided around the periphery of the hub disk. The shanks and mating slots are formed to have an interlocking geometry, such as a fir tree or dove tail interface, to prevent disengagement of the bladed pieces in a radial direction during high speed rotation of the rotor.
While enabling the production of a bladed rotor having a hub disk and bladed pieces fabricated from different alloys, the above-described inserted blade manufacturing approach is limited in certain respects. The formation of geometrically complex mating interfaces between the blade shanks and the hub disk can require multiple precision machining steps, which add undesired cost, duration, and complexity to the rotor manufacturing process. As a further drawback, the mating shank-disk interfaces can be difficult to seal. If not fully sealed, such interfaces can permit undesired leakage across the bladed rotor and trap debris increasing the propensity of the rotor to corrode. As a still further drawback, the formation of the mating shank-disk interfaces may necessitate an increase in the overall size and weight of the bladed rotor to achieve a structural integrity comparable to that of a monolithic or single piece bladed rotor. More recently, manufacturing approaches have been developed in which a full blade ring and a hub disk are separately produced from different superalloys and then metallurgically consolidated to produce a so-called “dual alloy bladed rotor.” As conventionally proposed and implemented, however, such approaches for manufacturing dual alloy bladed rotors remain limited in certain respects; e.g., such manufacturing approaches may prevent or render impractical the usage of single crystal and directionally-solidified superalloy in producing the rotor blades.
It is thus desirable to provide methods for manufacturing a dual alloy bladed rotor that overcome one or more of the above-described limitations. For example, it is desirable to provide methods for manufacturing dual alloy bladed rotors that reduce leakage across the bladed rotor, that minimize the overall size and weight of the bladed rotor, and that are generally less complex and costly to perform relative to other known rotor manufacturing approaches. Ideally, embodiments of such a manufacturing process would further enable fabrication of the rotor blades from a wide variety of high temperature materials including single crystal and directionally-solidified superalloys. It would also be desirable to provide embodiments of a dual alloy bladed rotor produced utilizing such a manufacturing process. Other desirable features and characteristics of embodiments of the present invention will become apparent from the subsequent Detailed Description and the appended Claims, taken in conjunction with the accompanying drawings and the foregoing Background.